The present invention relates to a blade/vane cascade segment, a blade/vane cascade, a stage, and a blade/vane channel of a turbomachine, as well as a turbomachine.
Turbomachines (such as gas and steam turbines) generally have a flow channel for conducting a fluid. The flow channel, which is also called an “annular space” is bounded radially inward by the shaft of a rotor and radially outward by a casing; the designations “radially” as well as “axially” and “peripheral direction”, and terms derived therefrom are always to be understood with reference to a (provided) axis of rotation of the rotor in this document—as long as nothing is indicated to the contrary.
Blade/vane cascades (for which the denotation “blade/vane ring” is also common) are arranged in the annular space of a turbomachine. They each comprise guide vanes or rotating blades that lie one behind the other in the peripheral direction at essentially regular distances, as well as stages belonging thereto, which are also called “cover plates”, and that have a stage edge on the inflow side and on the outflow side. These stage edges bound the stage surface facing the blades/vanes (or blade/vane elements) in the axial direction.
In this document, the stage edge “on the inflow side” is designated as the stage edge where the leading (axial) principal flow first passes through the annular space of the turbomachine during operation; correspondingly, the stage edge “on the outflow side” is the other edge. The principal flow thus passes through the stage bounded upstream/downstream by the stage edges on the inflow side/outflow side, i.e., the stage is directly adjacent to the flow. This stage, in particular, is thus not a “wing” displaced radially downward, which, during operation, is overlapped by an adjacent stage on the inflow side or the like, and is distanced from the principal flow that does not pass through it. The stage bounded by the stage edge on the inflow side and outflow side also does not comprise such a wing. The principal flow thus flows into the stage edge on the inflow side during operation. The indications “downstream” or “upstream”, respectively, refer correspondingly to the axial principal flow direction, and thus only to the axial position, regardless of a possible displacement in the peripheral direction: In particular, in this document, a point is to be understood as lying “downstream of the inflow edges” if it is arranged displaced axially in the direction of principal flow relative to a direct connection line between the inflow edges at the stage surface.
The pressure side of a blade/vane and the suction side of an adjacent blade/vane each bound a so-called blade/vane channel in the peripheral direction. In the radial direction, this blade/vane channel is bounded by so-called side walls within the turbomachine. These side walls are formed, on one hand, by the stages, and, on the other hand, by sections lying radially opposite to these stages: In the case of rotating blades, such a side wall is a radially outer-lying section (in particular, a section of the casing); in the case of guide vanes, it is a radially inner-lying section (in particular, a rotor hub).
The section of the stage surface that is bound in the axial direction by directly connecting the inflow (leading) edges or the outflow (trailing) edges, respectively, of adjacent blade/vane elements at the stage surface (or by a projection of a straight connection between the named edges in the radial direction onto the stage surface), and is bound in the peripheral direction by the suction side or pressure side thereof, is called in this document a “blade/vane intermediate strip”. The width of the blade/vane intermediate strip in the peripheral direction is named the “pitch distance” between the blade/vane cascade. It can be measured, in particular, as the distance between the leading edges of adjacent blades/vanes in the peripheral direction at the stage surface. The depth of the blade/vane intermediate space in the axial direction, thus the distance between the leading edges of the blade/vane elements and the trailing edges thereof that is measured parallel to the provided axis of rotation of the turbomachine is referred to as the “cascade span”.
A fluid flow conveyed through a flow channel is periodically influenced by the surfaces of the side walls. Flow layers that run next to these surfaces are more strongly diverted here, due to their slower speed, than flow layers that are further away from the side walls. Thus, a secondary flow that is superimposed on an axial principal flow arises and, in particular, leads to vortexes and pressure losses.
In order to reduce secondary flows, contouring is frequently introduced in the side walls in the form of elevations and/or depressions.
A plurality of these types of so-called “side wall contouring” are known from the prior art. By way of example, the patents or patent applications of the Applicant will be named: EP 2 487 329 B1; EP 2 787 172 A2; and EP 2 696 029 B1.
Furthermore, a flow channel having a side wall is known from the publication EP 1 126 132 A2, this channel having a radial depression in the region of the leading edges of the blade/vane elements. This depression extends in the axial direction over the majority of the flow channel and ends only just in front of, or perhaps behind the trailing edges. The surface of the flow-through region between leading and trailing edges will be locally enlarged thereby, which shall improve the efficiency of the rotor.
EP 2 372 088 A2 discloses an integrally fabricated turbine bladed disk, which has a ring with edges on the inflow and outflow sides, and rotating blades as well as depressions—in the region of the leading edges of the rotating blade elements—are arranged in the ring surface between these edges.